Geared Architecture for High Speed and Small Volume Fan Drive Turbine

ABSTRACT

A gas turbine engine includes a flex mount for a fan drive gear system. A very high speed fan drive turbine drives the fan drive gear system.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation of U.S. patent application Ser.No. 13/974,136, filed Aug. 23, 2013, which is a continuation of U.S.patent application Ser. No. 13/908,177, filed Jun. 3, 2013, which is acontinuation-in-part of U.S. patent application Ser. No. 13/623,309,filed Sep. 20, 2012, which is a continuation-in-part of U.S. patentapplication Ser. No. 13/342,508, filed Jan. 3, 2012, which claimspriority to U.S. Provisional Patent Application No. 61/494,453, filedJun. 8, 2011.

BACKGROUND

The present disclosure relates to a gas turbine engine, and moreparticularly to a flexible support structure for a geared architecturetherefor.

Epicyclic gearboxes with planetary or star ger trains may be used in gasturbine engines for their compact designs and efficient high gearreduction capabilities. Planetary and star gear trains generally includethree gear train elements: a central sun gear, an outer ring gear withinternal gear teeth, and a plurality of planet gears supported by aplanet carrier between and in meshed engagement with both the sun gearand the ring gear. The gear train elements share a common longitudinalcentral axis, about which at least two rotate. An advantage of epicyclicgear trains is that a rotary input can be connected to any one of thethree elements. One of the other two elements is then held stationarywith respect to the other two to permit the third to serve as an output.

In gas turbine engine applications, where a speed reduction transmissionis required, the central sun gear generally receives rotary input fromthe power plant, the outer ring gear is generally held stationary andthe planet gear carrier rotates in the same direction as the sun gear toprovide torque output at a reduced rotational speed. In star geartrains, the planet carrier is held stationary and the output shaft isdriven by the ring gear in a direction opposite that of the sun gear.

During flight, light weight structural cases deflect with aero andmaneuver loads causing significant amounts of transverse deflectioncommonly known as backbone bending of the engine. This deflection maycause the individual sun or planet gear's axis of rotation to loseparallelism with the central axis. This deflection may result in somemisalignment at gear train journal bearings and at the gear teeth mesh,which may lead to efficiency losses from the misalignment and potentialreduced life from increases in the concentrated stresses.

Further, with the geared architecture as set forth above, the torque andspeed of the input into the gear is quite high.

SUMMARY

In a featured embodiment, a gas turbine engine has a fan shaft driving afan, a frame supporting the fan shaft, and a plurality of gears to drivethe fan shaft. A flexible support at least partially supports theplurality of gears. The flexible support has a lesser stiffness than theframe. A first turbine section provides a drive input into the pluralityof gears. A second turbine section is also included. The first turbinesection has a first exit area at a first exit point and rotates at afirst speed. The second turbine section has a second exit area at asecond exit point and rotates at a second speed, which is faster thanthe first speed. A first performance quantity is defined as the productof the first speed squared and the first area. A second performancequantity is defined as the product of the second speed squared and thesecond area. A ratio of the first performance quantity to the secondperformance quantity is between about 0.5 and about 1.5.

In another embodiment according to the previous embodiment, the ratio isabove or equal to about 0.8.

In another embodiment according to any of the previous embodiments, thefirst turbine section has at least three stages.

In another embodiment according to any of the previous embodiments, thefirst turbine section has up to six stages.

In another embodiment according to any of the previous embodiments, thesecond turbine section has two or fewer stages.

In another embodiment according to any of the previous embodiments, apressure ratio across the first turbine section is greater than about5:1.

In another embodiment according to any of the previous embodiments, aratio of a thrust provided by the engine, to a volume of a turbinesection including both the high pressure turbine and the low pressureturbine is greater than or equal to about 1.5 and less than or equal toabout 5.5 lbf/inch³.

In another embodiment according to any of the previous embodiments, theframe includes a frame lateral stiffness and a frame transversestiffness. The flexible support includes a flexible support transversestiffness and a flexible support lateral stiffness. The flexible supportlateral stiffness is less than the frame lateral stiffness and theflexible support transverse stiffness is less than the frame transversestiffness.

In another embodiment according to any of the previous embodiments, aflexible coupling connects at least one of the plurality of gears to bedriven by the first turbine section.

In another embodiment according to any of the previous embodiments, theflexible coupling has a flexible coupling lateral stiffness and aflexible coupling transverse stiffness. The flexible coupling lateralstiffness is less than the frame lateral stiffness. The flexiblecoupling transverse stiffness is less than the frame transversestiffness.

In another embodiment according to any of the previous embodiments, theplurality of gears include a gear mesh that defines a gear mesh lateralstiffness and a gear mesh transverse stiffness. The gear mesh lateralstiffness is greater than the flexible support lateral stiffness. Thegear mesh transverse stiffness is greater than the flexible supporttransverse stiffness.

In another featured embodiment, a machine has a shaft driving a fan, aframe which supports the shaft, and a plurality of gears which drivesthe shaft. A flexible support, which at least partially supports theplurality of gears, has a lesser stiffness than the frame. A highpressure turbine and a low pressure turbine are included, the lowpressure turbine being configured to drive one of the plurality ofgears. A ratio of a thrust provided by the engine, to a volume of aturbine section including both the high pressure turbine and the lowpressure turbine, is are greater than or equal to about 1.5 and lessthan or equal to about 5.5 lbf/inch³.

In another embodiment according to the previous embodiment, the ratio isgreater than or equal to about 2.0.

In another embodiment according to any of the previous embodiments, theratio is greater than or equal to about 4.0.

In another embodiment according to any of the previous embodiments, thethrust is sea level take-off, flat-rated static thrust.

In another embodiment according to any of the previous embodiments, theframe includes a frame lateral stiffness and a frame transversestiffness. The flexible support includes a flexible support transversestiffness and a flexible support lateral stiffness. The flexible supportlateral stiffness is less than the frame lateral stiffness and theflexible support transverse stiffness is less than the frame transversestiffness.

In another embodiment according to any of the previous embodiments, aflexible coupling connects at least one of the plurality of gears to bedriven by the first turbine section.

In another embodiment according to any of the previous embodiments, theflexible coupling has a flexible coupling lateral stiffness and aflexible coupling transverse stiffness. The flexible coupling lateralstiffness is less than the frame lateral stiffness, and the flexiblecoupling transverse stiffness is less than the frame transversestiffness.

In another embodiment according to any of the previous embodiments, theplurality of gears include a gear mesh that defines a gear mesh lateralstiffness and a gear mesh transverse stiffness. The gear mesh lateralstiffness is greater than the flexible support lateral stiffness. Thegear mesh transverse stiffness is greater than the flexible supporttransverse stiffness.

In another featured embodiment, a gas turbine engine has a fan shaft anda frame which supports the fan shaft. The frame defines at least one ofa frame lateral stiffness and a frame transverse stiffness. A gearsystem drives the fan shaft. A flexible support at least partiallysupports the gear system. The flexible support defines at least one of aflexible support lateral stiffness with respect to the frame lateralstiffness and a flexible support transverse stiffness with respect tothe frame transverse stiffness. An input coupling to the gear systemdefines at least one of an input coupling lateral stiffness with respectto the frame lateral stiffness and an input coupling transversestiffness with respect to the frame transverse stiffness.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1A is a schematic cross-section of a gas turbine engine;

FIG. 1B shows a feature of the FIG. 1A engine.

FIG. 1C shows another feature.

FIG. 1D shows yet another feature.

FIG. 2 is an enlarged cross-section of a section of the gas turbineengine which illustrates a fan drive gear system (FDGS);

FIG. 3 is a schematic view of a flex mount arrangement for onenon-limiting embodiment of the FDGS;

FIG. 4 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of the FDGS;

FIG. 5 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a star system FDGS; and

FIG. 6 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a planetary system FDGS.

FIG. 7 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a star system FDGS; and

FIG. 8 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a planetary system FDGS.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. A mid-turbine frame 57 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion of the airflow passing therethrough.

The amount of thrust that can be produced by a particular turbinesection compared to how compact the turbine section is, is referred toas the power density, or the force density, of the turbine section, andis derived by the flat-rated Sea Level Take-Off (SLTO) thrust divided bythe volume of the entire turbine section. The example volume isdetermined from an inlet of the high pressure turbine 54 to an exit ofthe low pressure turbine 46. In order to increase the power density ofthe turbine section 28, each of the low pressure and high pressureturbines 46, 54 is made more compact. That is, the high pressure turbine54 and the low pressure turbine 46 are made with a shorter axial length,and the spacing between each of the turbines 46, 54 is decreased,thereby decreasing the volume of the turbine section 28.

The power density in the disclosed gas turbine engine 20 including thegear driven fan section 22 is greater than those provided in prior artgas turbine engine including a gear driven fan. Eight disclosedexemplary engines, which incorporate turbine sections and fan sectionsdriven through a reduction gear system and architectures as set forth inthis application, are described in Table I as follows:

TABLE 1 Turbine section Thrust/turbine Thrust SLTO volume from sectionvolume Engine (lbf) the Inlet (lbf/in³) 1 17,000 3,859 4.4 2 23,3005,330 4.37 3 29,500 6,745 4.37 4 33,000 6,745 4.84 5 96,500 31,086 3.1 696,500 62,172 1.55 7 96,500 46,629 2.07 8 37,098 6,745 5.50

In some embodiments, the power density is greater than or equal to about1.5 lbf/in³. In further embodiments, the power density is greater thanor equal to about 2.0 lbf/in³. In further embodiments, the power densityis greater than or equal to about 3.0 lbf/in³. In further embodiments,the power density is greater than or equal to about 4.0 lbf/in³. Infurther embodiments, the power density is less than or equal to about5.5 lbf/in³.

Engines made with the disclosed gear driven fan architecture, andincluding turbine sections as set forth in this application, providevery high efficiency operation, and increased fuel efficiency.

Referring to FIG. 1B, with continued reference to FIG. 1A, relativerotations between components of example disclosed engine architecture100 are schematically shown. In the example engine architecture 100, thefan 42 is connected, through the gearbox 48, to the low spool 30 towhich the low pressure compressor 44 and the low pressure turbine 46 areconnected. The high pressure compressor 52 and the high pressure turbine54 are connected to a common shaft forming the high spool 32. The highspool 32 rotates opposite the direction of rotation of the fan 42(illustrated in FIG. 1B as the “+” direction.) The low spool 30 rotatesin the same direction as the fan 42 (illustrated in FIG. 1B as the “−”direction.) The high pressure turbine 54 and the low pressure turbine46, along with the mid-turbine frame 57 together forms the turbinesection 28 of the gas turbine engine 20. Other relative rotationdirections between the two spools and the fan come within the scope ofthis disclosure.

One disclosed example speed change device 48 has a gear reduction ratioexceeding 2.3:1, meaning that the low pressure turbine 46 turns at least2.3 times faster than the fan 42. An example disclosed speed changedevice is an epicyclical gearbox of a planet type, where the input is tothe center “sun” gear 260. Planet gears 262 (only one shown) around thesun gear 260 rotate and are spaced apart by a carrier 264 that rotatesin a direction common to the sun gear 260. A ring gear 266, which isnon-rotatably fixed to the engine static casing 36 (shown in FIG. 1),contains the entire gear assembly. The fan 42 is attached to and drivenby the carrier 264 such that the direction of rotation of the fan 42 isthe same as the direction of rotation of the carrier 264 that, in turn,is the same as the direction of rotation of the input sun gear 260.Accordingly, the low pressure compressor 44 and the low pressure turbine46 counter-rotate relative to the high pressure compressor 52 and thehigh pressure turbine 54.

Counter rotating the low pressure compressor 44 and the low pressureturbine 46 relative to the high pressure compressor 52 and the highpressure turbine 54 provides certain efficient aerodynamic conditions inthe turbine section 28 as the generated high speed exhaust gas flowmoves from the high pressure turbine 54 to the low pressure turbine 46.Moreover, the mid-turbine frame 57 contributes to the overallcompactness of the turbine section 28. Further, the airfoil 59 of themid-turbine frame 57 surrounds internal bearing support structures andoil tubes that are cooled. The airfoil 59 also directs flow around theinternal bearing support structures and oil tubes for streamlining thehigh speed exhaust gas flow. Additionally, the airfoil 59 directs flowexiting the high pressure turbine 54 to a proper angle desired topromote increased efficiency of the low pressure turbine 46.

Flow exiting the high pressure turbine 54 has a significant component oftangential swirl. The flow direction exiting the high pressure turbine54 is set almost ideally for the blades in a first stage of the lowpressure turbine 46 for a wide range of engine power settings. Thus, theaerodynamic turning function of the mid turbine frame 57 can beefficiently achieved without dramatic additional alignment of airflowexiting the high pressure turbine 54.

Referring to FIG. 1C, the example turbine section 28 volume isschematically shown and includes first, second and third stages 46A, 46Band 46C. Each of the stages 46A, 46B and 46C includes a correspondingplurality of blades 212 and vanes 214. The example turbine sectionfurther includes an example air-turning vane 220 between the low andhigh turbines 54, 46 that has a modest camber to provide a small degreeof redirection and achieve a desired flow angle relative to blades 212of the first stage 46 a of the low pressure turbine 46. The disclosedvane 220 could not efficiently perform the desired airflow function ifthe low and high pressure turbines 54, 46 rotated in a common direction.

The example mid-turbine frame 57 includes multiple air turning vanes 220in a row that direct air flow exiting the high pressure turbine 54 andensure that air is flowing in the proper direction and with the properamount of swirl. Because the disclosed turbine section 28 is morecompact than previously utilized turbine sections, air has less distanceto travel between exiting the mid-turbine frame 57 and entering the lowpressure turbine 46. The smaller axial travel distance results in adecrease in the amount of swirl lost by the airflow during thetransition from the mid-turbine frame 57 to the low pressure turbine 46,and allows the vanes 220 of the mid-turbine frame 57 to function asinlet guide vanes of the low pressure turbine 46. The mid-turbine frame57 also includes a strut 221 providing structural support to both themid-turbine frame 57 and to the engine housing. In one example, themid-turbine frame 57 is much more compact by encasing the strut 221within the vane 220, thereby decreasing the length of the mid-turbineframe 57.

At a given fan tip speed and thrust level provided by a given fan size,the inclusion of the speed change device 48 (shown in FIGS. 1A and 1B)provides a gear reduction ratio, and thus the speed of the low pressureturbine 46 and low pressure compressor 44 components may be increased.More specifically, for a given fan diameter and fan tip speed, increasesin gear ratios provide for a faster turning turbine that, in turn,provides for an increasingly compact turbine and increased thrust tovolume ratios of the turbine section 28. By increasing the gearreduction ratio, the speed at which the low pressure compressor 44 andthe low pressure turbine 46 turn, relative to the speed of the fan 42,is increased.

Increases in rotational speeds of the gas turbine engine 20 componentsincreases overall efficiency, thereby providing for reductions in thediameter and the number of stages of the low pressure compressor 44 andthe low pressure turbine 46 that would otherwise be required to maintaindesired flow characteristics of the air flowing through the core flowpath C. The axial length of each of the low pressure compressor 44 andthe low pressure turbine 46 can therefore be further reduced due toefficiencies gained from increased speed provided by an increased gearratio. Moreover, the reduction in the diameter and the stage count ofthe turbine section 28 increases the compactness and provides for anoverall decrease in required axial length of the example gas turbineengine 20.

In order to further improve the thrust density of the gas turbine engine20, the example turbine section 28 (including the high pressure turbine54, the mid-turbine frame 57, and the low pressure turbine 46) is mademore compact than traditional turbine engine designs, thereby decreasingthe length of the turbine section 28 and the overall length of the gasturbine engine 20.

In order to make the example low pressure turbine 46 compact, make thediameter of the low pressure turbine 46 more compatible with the highpressure turbine 54, and thereby make the air-turning vane 220 of themid-turbine frame 57 practical, stronger materials in the initial stagesof the low pressure turbine 46 may be required. The speeds andcentrifugal pull generated at the compact diameter of the low pressureturbine 46 pose a challenge to materials used in prior art low pressureturbines.

Examples of materials and processes within the contemplation of thisdisclosure for the air-turning vane 220, the low pressure turbine blades212, and the vanes 214 include materials with directionally solidifiedgrains to provided added strength in a span-wise direction. An examplemethod for creating a vane 220, 214 or turbine blade 212 havingdirectionally solidified grains can be found in U.S. application Ser.No. 13/290667, and U.S. Pat. Nos. 7,338,259 and 7,871,247, each of whichis incorporated by reference. A further, engine embodiment utilizes acast, hollow blade 212 or vane 214 with cooling air introduced at theleading edge of the blade/vane and a trailing edge discharge of thecooling air. Another embodiment uses an internally cooled blade 212 orvane 214 with film cooling holes. An additional engine embodimentutilizes an aluminum lithium material for construction of a portion ofthe low pressure turbine 46. The example low pressure turbine 46 mayalso be constructed utilizing at a powdered metal disc or rotor.

Additionally, one or more rows of turbine blades 212 of the low pressureturbine 46 can be constructed using a single crystal blade material.Single crystal constructions oxidize at higher temperatures as comparedto non-single crystal constructions and thus can withstand highertemperature airflow. Higher temperature capability of the turbine blades212 provide for a more efficient low pressure turbine 46 that may befurther reduced in size.

While the illustrated low pressure turbine 46 includes three turbinestages 46 a, 46 b, and 46 c, the low pressure turbine 46 can be modifiedto include up to six turbine stages. Increasing the number of lowpressure turbine stages 46 a, 46 b, 46 c at constant thrust slightlyreduces the thrust density of the turbine section 28 but also increasespower available to drive the low pressure compressor and the fan section22.

Further, the example turbine blades may be internally cooled to allowthe material to retain a desired strength at higher temperatures andthereby perform as desired in view of the increased centrifugal forcegenerated by the compact configuration while also withstanding thehigher temperatures created by adding low pressure compressor 44 stagesand increasing fan tip diameter.

Each of the disclosed embodiments enables the low pressure turbine 46 tobe more compact and efficient, while also improving radial alignment tothe high pressure turbine 54. Improved radial alignment between the lowand high pressure turbines 54, 46 increases efficiencies that can offsetany increases in manufacturing costs incurred by including the airturning vane 220 of the mid-turbine frame 57.

In light of the foregoing embodiments, the overall size of the turbinesection 28 has been greatly reduced, thereby enhancing the engine'spower density. Further, as a result of the improvement in power density,the engine's overall propulsive efficiency has been improved.

An exit area 400 is shown, in FIG. 1D and FIG. 1A, at the exit locationfor the high pressure turbine section 54. An exit area for the lowpressure turbine section is defined at exit 401 for the low pressureturbine section. As shown in FIG. 1D, the turbine engine 20 may becounter-rotating. This means that the low pressure turbine section 46and low pressure compressor section 44 rotate in one direction, whilethe high pressure spool 32, including high pressure turbine section 54and high pressure compressor section 52 rotate in an opposed direction.The gear reduction 48, which may be, for example, an epicyclictransmission (e.g., with a sun, ring, and star gears), is selected suchthat the fan 42 rotates in the same direction as the high spool 32. Withthis arrangement, and with the other structure as set forth above,including the various quantities and operational ranges, a very highspeed can be provided to the low pressure spool. Low pressure turbinesection and high pressure turbine section operation are often evaluatedlooking at a performance quantity which is the exit area for the turbinesection multiplied by its respective speed squared. This performancequantity (“PQ”) is defined as:

PQ_(ltp)=(A _(lpt) ×V _(lpt) ²)  Equation 1:

PQ_(hpt)=(A _(hpt) ×V _(hpt) ²)  Equation 2:

where A_(lpt) is the area of the low pressure turbine section at theexit thereof (e.g., at 401), where V_(lpt) is the speed of the lowpressure turbine section, where A_(hpt) is the area of the high pressureturbine section at the exit thereof (e.g., at 400), and where V_(hpt) isthe speed of the low pressure turbine section.

Thus, a ratio of the performance quantity for the low pressure turbinesection compared to the performance quantify for the high pressureturbine section is:

(A _(lpt) ×V _(lpt) ²)/(A _(hpt) ×V _(hpt)²)=PQ_(ltp/)PQ_(hpt)  Equation 3:

In one turbine embodiment made according to the above design, the areasof the low and high pressure turbine sections are 557.9 in² and 90.67in², respectively. Further, the speeds of the low and high pressureturbine sections are 10179 rpm and 24346 rpm, respectively. Thus, usingEquations 1 and 2 above, the performance quantities for the low and highpressure turbine sections are:

PQ_(lpt)=(A _(lpt) ×V _(lpt) ²)=(557.9 in²)(10179 rpm)²=57805157673.9in² rpm²  Equation 1:

PQ_(hpt)=(A _(hpt) ×V _(hpt) ²)=(90.67 in²)(24346 rpm)²=53742622009.72in² rpm²  Equation 2:

and using Equation 3 above, the ratio for the low pressure turbinesection to the high pressure turbine section is:

Ratio=PQ_(lpt/)PQ_(hpt)=57805 157673.9 in² rpm²/53742622009.72 in²rpm²=1.075

In another embodiment, the ratio was about 0.5 and in another embodimentthe ratio was about 1.5. With PQ_(ltp/) PQ_(hpt) ratios in the 0.5 to1.5 range, a very efficient overall gas turbine engine is achieved. Morenarrowly, PQ_(ltp/) PQ_(hpt) ratios of above or equal to about 0.8 aremore efficient. Even more narrowly, PQ_(lpt/) PQ_(hpt) ratios above orequal to 1.0 are even more efficient. As a result of these PQ_(ltp/)PQ_(hpt) ratios, in particular, the turbine section can be made muchsmaller than in the prior art, both in diameter and axial length. Inaddition, the efficiency of the overall engine is greatly increased.

The low pressure compressor section is also improved with thisarrangement, and behaves more like a high pressure compressor sectionthan a traditional low pressure compressor section. It is more efficientthan the prior art, and can provide more work in fewer stages. The lowpressure compressor section may be made smaller in radius and shorter inlength while contributing more toward achieving the overall pressureratio design target of the engine.

A worker of ordinary skill in the art, being apprised of the disclosureabove, would recognize that high torque and high speed will be presentedby the low speed spool 30 into the gear architecture 48. Thus, aflexible mount arrangement becomes important.

With reference to FIG. 2, the geared architecture 48 generally includesa fan drive gear system (FDGS) 60 driven by the low speed spool 30(illustrated schematically) through an input coupling 62. The inputcoupling 62 both transfers torque from the low speed spool 30 to thegeared architecture 48 and facilitates the segregation of vibrations andother transients therebetween. In the disclosed non-limiting embodiment,the FDGS 60 may include an epicyclic gear system which may be, forexample, a star system or a planet system.

The input coupling 62 may include an interface spline 64 joined, by agear spline 66, to a sun gear 68 of the FDGS 60. The sun gear 68 is inmeshed engagement with multiple planet gears 70, of which theillustrated planet gear 70 is representative. Each planet gear 70 isrotatably mounted in a planet carrier 72 by a respective planet journalbearing 75. Rotary motion of the sun gear 68 urges each planet gear 70to rotate about a respective longitudinal axis P. The gears may begenerally as shown schematically in FIG. 1B.

Each planet gear 70 is also in meshed engagement with rotating ring gear74 that is mechanically connected to a fan shaft 76. Since the planetgears 70 mesh with both the rotating ring gear 74 as well as therotating sun gear 68, the planet gears 70 rotate about their own axes todrive the ring gear 74 to rotate about engine axis A. The rotation ofthe ring gear 74 is conveyed to the fan 42 (FIG. 1) through the fanshaft 76 to thereby drive the fan 42 at a lower speed than the low speedspool 30. It should be understood that the described geared architecture48 is but a single non-limiting embodiment and that various other gearedarchitectures will alternatively benefit herefrom.

With reference to FIG. 3, a flexible support 78 supports the planetcarrier 72 to at least partially support the FDGS 60A with respect tothe static structure 36 such as a front center body which facilitatesthe segregation of vibrations and other transients therebetween. Itshould be understood that various gas turbine engine case structures mayalternatively or additionally provide the static structure and flexiblesupport 78. It is to be understood that the term “lateral” as usedherein refers to a perpendicular direction with respect to the axis ofrotation A and the term “transverse” refers to a pivotal bendingmovement with respect to the axis of rotation A so as to absorbdeflections which may be otherwise applied to the FDGS 60. The staticstructure 36 may further include a number 1 and 1.5 bearing supportstatic structure 82 which is commonly referred to as a “K-frame” whichsupports the number 1 and number 1.5 bearing systems 38A. 38B. Notably,the K-frame bearing support defines a lateral stiffness (represented asKframe in FIG. 3) and a transverse stiffness (represented asKframe^(BEND) in FIG. 3) as the referenced factors in this non-limitingembodiment.

In this disclosed non-limiting embodiment, the lateral stiffness (KFS;KIC) of both the flexible support 78 and the input coupling 62 are eachless than about 11% of the lateral stiffness (Kframe). That is, thelateral stiffness of the entire FDGS 60 is controlled by this lateralstiffness relationship. Alternatively, or in addition to thisrelationship, the transverse stiffness of both the flexible support 78and the input coupling 62 are each less than about 11% of the transversestiffness (Kframe^(BEND)). That is, the transverse stiffness of theentire FDGS 60 is controlled by this transverse stiffness relationship.

With reference to FIG. 4, another non-limiting embodiment of a FDGS 60Bincludes a flexible support 78′ that supports a rotationally fixed ringgear 74′. The fan shaft 76′ is driven by the planet carrier 72′ in theschematically illustrated planet system which otherwise generallyfollows the star system architecture of FIG. 3.

With reference to FIG. 5, the lateral stiffness relationship within aFDGS 60C itself (for a star system architecture) is schematicallyrepresented. The lateral stiffness (KIC) of an input coupling 62, alateral stiffness (KFS) of a flexible support 78, a lateral stiffness(KRG) of a ring gear 74 and a lateral stiffness (KJB) of a planetjournal bearing 75 are controlled with respect to a lateral stiffness(KGM) of a gear mesh within the FDGS 60.

In the disclosed non-limiting embodiment, the stiffness (KGM) may bedefined by the gear mesh between the sun gear 68 and the multiple planetgears 70. The lateral stiffness (KGM) within the FDGS 60 is thereferenced factor and the static structure 82′ rigidly supports the fanshaft 76. That is, the fan shaft 76 is supported upon bearing systems38A, 38B which are essentially rigidly supported by the static structure82′. The lateral stiffness (KJB) may be mechanically defined by, forexample, the stiffness within the planet journal bearing 75 and thelateral stiffness (KRG) of the ring gear 74 may be mechanically definedby, for example, the geometry of the ring gear wings 74L, 74R (FIG. 2).

In the disclosed non-limiting embodiment, the lateral stiffness (KRG) ofthe ring gear 74 is less than about 12% of the lateral stiffness (KGM)of the gear mesh; the lateral stiffness (KFS) of the flexible support 78is less than about 8% of the lateral stiffness (KGM) of the gear mesh;the lateral stiffness (KJB) of the planet journal bearing 75 is lessthan or equal to the lateral stiffness (KGM) of the gear mesh; and thelateral stiffness (KIC) of an input coupling 62 is less than about 5% ofthe lateral stiffness (KGM) of the gear mesh.

With reference to FIG. 6, another non-limiting embodiment of a lateralstiffness relationship within a FDGS 60D itself are schematicallyillustrated for a planetary gear system architecture, which otherwisegenerally follows the star system architecture of FIG. 5.

It should be understood that combinations of the above lateral stiffnessrelationships may be utilized as well. The lateral stiffness of each ofstructural components may be readily measured as compared to filmstiffness and spline stiffness which may be relatively difficult todetermine.

By flex mounting to accommodate misalignment of the shafts under designloads, the FDGS design loads have been reduced by more than 17% whichreduces overall engine weight. The flex mount facilitates alignment toincrease system life and reliability. The lateral flexibility in theflexible support and input coupling allows the FDGS to essentially‘float’ with the fan shaft during maneuvers. This allows: (a) the torquetransmissions in the fan shaft, the input coupling and the flexiblesupport to remain constant during maneuvers; (b) maneuver inducedlateral loads in the fan shaft (which may otherwise potentially misaligngears and damage teeth) to be mainly reacted to through the number 1 and1.5 bearing support K-frame; and (c) both the flexible support and theinput coupling to transmit small amounts of lateral loads into the FDGS.The splines, gear tooth stiffness, journal bearings, and ring gearligaments are specifically designed to minimize gear tooth stressvariations during maneuvers. The other connections to the FDGS areflexible mounts (turbine coupling, case flex mount). These mount springrates have been determined from analysis and proven in rig and flighttesting to isolate the gears from engine maneuver loads. In addition,the planet journal bearing spring rate may also be controlled to supportsystem flexibility.

FIG. 7 is similar to FIG. 5 but shows the transverse stiffnessrelationships within the FDGS 60C (for a star system architecture). Thetransverse stiffness (KIC^(BEND)) of the input coupling 62, a transversestiffness (KFS^(BEND)) of the flexible support 78, a transversestiffness (KRG^(BEND)) of the ring gear 74 and a transverse stiffness(KJB^(BEND)) of the planet journal bearing 75 are controlled withrespect to a transverse stiffness (KGM^(BEND)) of the gear mesh withinthe FDGS 60.

In the disclosed non-limiting embodiment, the stiffness (KGM^(BEND)) maybe defined by the gear mesh between the sun gear 68 and the multipleplanet gears 70. The transverse stiffness(KGM^(BEND)) within the FDGS 60is the referenced factor and the static structure 82′ rigidly supportsthe fan shaft 76. That is, the fan shaft 76 is supported upon bearingsystems 38A, 38B which are essentially rigidly supported by the staticstructure 82′. The transverse stiffness (KJB^(BEND)) may be mechanicallydefined by, for example, the stiffness within the planet journal bearing75 and the transverse stiffness (KRG^(BEND)) of the ring gear 74 may bemechanically defined by, for example, the geometry of the ring gearwings 74L, 74R (FIG. 2).

In the disclosed non-limiting embodiment, the transverse stiffness(KRG^(BEND)) of the ring gear 74 is less than about 12% of thetransverse stiffness (KGM^(BEND)) of the gear mesh; the transversestiffness (KFS^(BEND)) of the flexible support 78 is less than about 8%of the transverse stiffness (KGM^(BEND)) of the gear mesh; thetransverse stiffness (KJB^(BEND)) of the planet journal bearing 75 isless than or equal to the transverse stiffness (KGM^(BEND)) of the gearmesh; and the transverse stiffness (KIC^(BEND)) of an input coupling 62is less than about 5% of the transverse stiffness (KGM^(BEND)) of thegear mesh.

FIG. 8 is similar to FIG. 6 but shows the transverse stiffnessrelationship within the FDGS 60D for the planetary gear systemarchitecture.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The combined arrangement of the high power density and fan drive turbinewith the high AN² performance quantity, all incorporated with theflexible mounting structure, provide a very robust and efficient gasturbine engine.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A gas turbine engine comprising: a fan shaftdriving a fan; a frame which supports said fan shaft; a plurality ofgears to drive said fan shaft; a flexible support which at leastpartially supports said plurality of gears, said flexible support havinga different stiffness than said frame; a first turbine section providinga drive input into said plurality of gears; and a second turbinesection, wherein said first turbine section has a first exit area at afirst exit point and rotates at a first speed, wherein said secondturbine section has a second exit area at a second exit point androtates at a second speed, which is faster than the first speed, whereina first performance quantity is defined as the product of the firstspeed squared and the first area, wherein a second performance quantityis defined as the product of the second speed squared and the secondarea, and wherein a ratio of the first performance quantity to thesecond performance quantity is between about 0.5 and about 1.5.
 2. Theturbine section as set forth in claim 1, wherein said ratio is above orequal to about 0.8.
 3. The turbine section as set forth in claim 1,wherein said first turbine section has at least 3 stages.
 4. The turbinesection as set forth in claim 1, wherein said first turbine section hasup to 6 stages.
 5. The turbine section as set forth in claim 1, whereinsaid second turbine section has 2 or fewer stages.
 6. The turbinesection as set forth in claim 1, wherein a pressure ratio across thefirst turbine section is greater than about 5:1.
 7. The gas turbineengine as set forth in claim 1, including a ratio of a thrust providedby said engine, to a volume of a turbine section including both saidhigh pressure turbine and said low pressure turbine being greater thanor equal to about 1.5 and less than or equal to about 5.5 lbf/inch³. 8.The gas turbine engine as set forth in claim 1, wherein said frameincludes a frame lateral stiffness and a frame transverse stiffness, andsaid flexible support includes a flexible support transverse stiffnessand a flexible support lateral stiffness, and said flexible supportlateral stiffness being less than said frame lateral stiffness and saidflexible support transverse stiffness being less than said frametransverse stiffness.
 9. The gas turbine engine as set forth in claim 8,wherein a flexible coupling connects at least one of said plurality ofgears to be driven by said first turbine section.
 10. The gas turbineengine as set forth in claim 9, wherein said flexible coupling has aflexible coupling lateral stiffness and a flexible coupling transversestiffness, and said flexible coupling lateral stiffness being less thansaid frame lateral stiffness, and said flexible coupling transversestiffness being less than said frame transverse stiffness.
 11. The gasturbine engine as set forth in claim 8, wherein said plurality of gearsinclude a gear mesh that defines a gear mesh lateral stiffness and agear mesh transverse stiffness, said gear mesh lateral stiffness beinggreater than said flexible support lateral stiffness and said gear meshtransverse stiffness being greater than said flexible support transversestiffness.
 12. A machine comprising: a shaft driving a fan; a framewhich supports said shaft; a plurality of gears which drives the shaft;a flexible support which at least partially supports said plurality ofgears, said flexible support having a different stiffness than saidframe; a high pressure turbine; a low pressure turbine, said lowpressure turbine being configured to drive one of said plurality ofgears; a ratio of a thrust provided by said engine, to a volume of aturbine section including both said high pressure turbine and said lowpressure turbine being greater than or equal to about 1.5 and less thanor equal to about 5.5 lbf/inch³.
 13. The gas turbine engine as set forthin claim 12, wherein said ratio is greater than or equal to about 2.0.14. The gas turbine engine as set forth in claim 13, wherein said ratiois greater than or equal to about 4.0.
 15. The gas turbine engine as setforth in claim 12, wherein said thrust is sea level take-off, flat-ratedstatic thrust.
 16. The gas turbine engine as set forth in claim 12,wherein said frame includes a frame lateral stiffness and a frametransverse stiffness, and said flexible support includes a flexiblesupport transverse stiffness and a flexible support lateral stiffness,and said flexible support lateral stiffness being less than said framelateral stiffness and said flexible support transverse stiffness beingless than said frame transverse stiffness.
 17. The gas turbine engine asset forth in claim 16, wherein a flexible coupling connects at least oneof said plurality of gears to be driven by said first turbine section.18. The gas turbine engine as set forth in claim 17, wherein saidflexible coupling has a flexible coupling lateral stiffness and aflexible coupling transverse stiffness, and said flexible couplinglateral stiffness being less than said frame lateral stiffness, and saidflexible coupling transverse stiffness being less than said frametransverse stiffness.
 19. The gas turbine engine as set forth in claim16, wherein said plurality of gears include a gear mesh that defines agear mesh lateral stiffness and a gear mesh transverse stiffness, saidgear mesh lateral stiffness being greater than said flexible supportlateral stiffness and said gear mesh transverse stiffness being greaterthan said flexible support transverse stiffness.
 20. A gas turbineengine comprising: a fan shaft; a frame which supports said fan shaft,said frame defines a frame lateral stiffness and a frame transversestiffness; a gear system which drives said fan shaft; a flexible supportwhich at least partially supports said gear system, said flexiblesupport defines a flexible support lateral stiffness with respect tosaid frame lateral stiffness and a flexible support transverse stiffnesswith respect to said frame transverse stiffness; and an input couplingto said gear system, said input coupling defines an input couplinglateral stiffness with respect to said frame lateral stiffness and aninput coupling transverse stiffness with respect to said frametransverse stiffness.